Guided missile control systems

ABSTRACT

1. In an antiaircraft system, target tracking means for continuously establishing the position of a target aircraft in space, missile tracking means for similarly establishing the position of a missile, computing means responsive to position information from both said target and missile tracking means to produce control signals for said missile to insure interception of the target thereby and means for substituting the target tracking means for said missile tracking means as a source of missile position information for said computing means during the final phases of an engagement.

United States Patent Schaefer May 13, 1975 GUIDED MISSILE CONTROLSYSTEMS Primary Examiner-Benjamin A. Borchelt Assistant Examiner-ThomasH. Webb b W. S haefer Watchun NJ. [75] Inventor Jaco c g Attorney,Agent, or Firm-E. W. Adams, Jr. [73] Assignee: Bell TelephoneLaboratories.

Incorporated, Murray Hill, NJ. EXEMPLARY CLAIM [22] Filed: July 30,1956 1. In an antiaircraft system target tracking means for continuouslyestablishing the position of a target air- [21] Appl' 600858 craft inspace, missile tracking means for similarly establishing the position ofa missile, computing means [52] U.S. Cl 244/3.l3; 244/3.l9 responsive toposition information from both said tar- [51] Int. Cl F4lg 9/00; F4lg 1H00; F4lg 7/]4 get and missile tracking means to produce control sig-[58] Field of Search 244/14, 3.13, 3.l4 nals for said missile to insureinterception of the target thereby and means for substituting the targettracking [56] References Cited means for said missile tracking means asa source of UNITED STATES PATENTS missile position information for saidcomputing means 2 801 815 8/1957 Williams 244/14 during the final phasesof an engagement" 7 Claims, 4 Drawing Figures TARGET /'/X u- PRED/CTEDPOS/T/OIVAT INTERCEPT MISS/LE7 /6 25 26 [02 U6 0 a fix I /a4 l TRANS. c.,yggg egg-7, R c. m 5, [06

RADAR RADAR 23%? L Z fi/"T 5 k M M 4T5) COMPUTER 2a [4 //o L -29 AMEMauRsroRoER (T=o) TIME OF FLIGHT YAWORDER 0/v/0ER 6,4 Mus/LE PATH (PITCHoRoER igggfifl gg gig? D D R 55, 5' fx v hm Ros.

& ERRoR BAiL/sT/c 32 COMM MISS/LE Cl-IA Ann I g g $333 ls i'lcs um HM rr r f MISS/LE 3/ 30 HEAD/N6 I RESOtL vER 22 I PRED/CTER //z\.|CORRECTION FRED/67' L L ":5? {at-001502 //4 l CONTROL flea CIRCUIT WEE?2 OF d PATENTEQ kiAY 7 3 $75 INVENTOR By J. SCHAEFER ATTORNEY PATENIEB WI 31875 SHEET 3 OF 4 IN I/EN TOR B M. SCHA EFER A T TOR/V5 Y GUIDEDMISSILE CONTROL SYSTEMS This invention relates to guided missile systemsand more particularly to improvements in those missile systems employingcommand guidance techniques.

In a copending application, Ser. No. 449,396, filed Aug. 12, 1954 in thenames of E. L. Norton and the present inventor and assigned to theassignee of the present application, which matured into US. Pat. No.3,156,435 on Nov. I0, 1964. there was disclosed a guided missile systemdesigned for the interception and destruction of high altitude, highperformance bombing aircraft against which adequate defense withconventional antiaircraft artillery is inoperative. The extreme rangesat which engagement must occur to prevent a successful bombing attackmake it necessary to provide some means for adjusting the course of aprojectile launched against the bomber during the rather considerabletime of flight between launching and interception.

As disclosed in the above-mentioned copending application, a practicalguided missile system for such ap plications may include a missile,normally selfpropelled, which may be controlled in accordance withcommands issued by a ground based guidance equipment. This guidanceequipment includes precision radars for individually tracking both thetarget and the missile launched against the target to obtain data as tothe present positions of both. These data are supplied to a computerwhich predicts the future position of the target at an assumed time ofinterception and generates orders for transmission to the missile duringits flight to control the course thereof in such a way as to insureinterception of the target at the assumed time, Equipment aboard themissile provides a frame of reference traveling with the missile andidentifiable at the location of the ground guidance equipment withrespect to which control orders may be produced.

Although the command guidance system is considered highly efficientthere are other systems of missile guidance which have attractivefeatures. One of these which is now well known is the co-called beamriding system of missile guidance. Here a tracking device, usually aground based radar is employed to track the target continuously and themissile is launched in such a way as to intercept the beam of the radarshortly after launch. Detection devices aboard the missile respond tothe beam when intercepted and provide inputs for a missile-bornecomputer which determines those adjustments in the controls of themissile which are required to maintain the missile in the beam as thebeam is swept to follow the flight of the target. This system has theinherent advantage that a single radar or guidance device is employed.In more detail the use of a single radar eliminates the need for precisecorrections for parallax between the locations of the two trackingradars of the command system. In addition, since only a single radar isemployed bore-sighting errors are not compounded. Normally such mattersdo not constitute so great a disadvantage as to outweigh the greateradvantages of the command system. However, at extreme ranges and duringthe final phases of an engagement slight resultant misalignments betweenthe two radars may seriously affect performance. On the other hand thebeam riding system has the serious disadvantage that the missile doesnot follow the most efficient course (in terms of time and fuel) to thetarget.

It is an object of the present invention to so modify a command guidancesystem as to make available therein the attractive features of the beamriding missile system of guidance without incurring the penaltiesinherent in the beam riding system.

In accordance with this object the command missile system of the presentinvention is generally the same as that of the copending applicationreferred to above. In addition to the elements thereof as outlinedabove, means are provided for determining when the beams of the trackingradars for both missile and target have approached one another to withinpredetermined limits. At this time, which indicates effective proximityof the missile and the target, means are actuated for transferring themissile tracking function from the missile tracking radar to the targettracking radar. For this purpose modifications are made in the targettracking radar which enable it to receive both return pulses from thetarget and those from the missile which occur at a different frequency.The switching device referred to above transfers the missile positioninputs to the computer from the missile tracking radar to the targettracking radar and thus makes available for the final phases of anattack the advantageous single reference system of guidance normallyassociated with a beam riding system.

The above and other features of the invention will be described indetail in the following specification taken in connection with thedrawings, in which FIG. 1 is a block schematic diagram of the completecommand missile guidance system of the invention;

FIG. 2 is a diagram in schematic form of the missile employed in thesystem of FIG. 1 showing the equipment required aboard the missile;

FIG. 3 is a vector diagram illustrating the missile command problem andthe manner in which the orders for transmission to the missile arecomputed; and

FIG. 4 is a block diagram of the control system required fortransferring the missile tracking function from the missile trackingradar of the target tracking radar and provides details of the controlcircuit shown schematically in FIG. 1.

In the broadest sense the antiaircraft guided missile system of theinvention comprises a target tracking device 10, a missile trackingdevice 12, a computer 14 and a missile 16, all as shown in the blockdiagram of FIG. 1. Target tracking device 10 which preferably comprisesa precision automatic tracking radar is ranged continuously to providedata as to the present position of a target aircraft 18. This trackingradar may, for example, be similar to the well known SCR-584 radar whichis described in detail in Electronics" for November 1945 beginning atpage 104, for December 1945 beginning at page 104 and for February 1946beginning at page 1 10. Briefly this radar is an automatic trackingradar employing conical lobing whereby azimuth and elevation errorsignals are produced by the receiver. These signals may be applied toservo systems to cause the radar antenna to continuously track thetarget in both elevation and azimuth. In addition this radar includes arange unitalso responsive to the reflected radar pulses whichautomatically maintains itself adjusted to represent the slant range tothe target.

As shown in FIG. 1 target tracking radar 10 includes a transmitteroperating at a frequency f which is used only by this radar in thepresent system, a receiver unit 102 tuned to receive echo signals at afrequency f and a second receiver unit 104 also associated with the sameradar antenna and tuned to receive echoes at a frequency f; which aswill be described below identifies those return signals emanating fromthe missile. Receiver units 102 and 104 may be identical with theexception of the frequency to which they are tuned and may convenientlybe connected to the same radio frequency components of the radartransmit-receive equipment. This can be accomplished at radio frequencythrough the use of a conventional branching network, the output of whichis applied to two preamplifiers each associated with its own mixer,intermediate frequency amplifier, and other necessary receivercircuitry. As will appear below, receiver unit 104 need not include arange unit.

The normal output of tracking radar is derived from informationindicating the elevation and azimuth errors and the range to the targetas indicated by the error signals received by receiver 102 andtranslated into position data appropriate for transmission to thecomputer in the range unit forming a part of the receiver and in servounit 106 which derives the azimuth and elevation signals while orientingthe antenna to reduce the error signals to zero. This servo unit may beconsidered to be a typical unit and is assumed for the present purposesthat the azimuth and elevation error signals and the range signals areconverted to electrical quantities corresponding to the azimuth rangeand elevation for individual transmission over a connector to apredictor 22. The form in which the target position data are determinedis not significant since the means for converting data in one coordinatesystem to another system are well known in the computer art.

Target radar 10 produces a second set of output quantities forapplication to transmission link 108 which quantities are of the samenature as those applied to transmission link 20 and are representativeof azimuth and elevation data. In this instance, however, they arederived from receiver 104 tuned to the frequency f and converted in dataunit 110. It should by noted that the quantities obtained from data unit110 are employed only for application to transmission link 108 and arenot employed in the automatic tracking circuitry of target trackingradar 10.

The missile tracking device 12 may be similar to target tracking device10 and may in the same way produce output quantities proportional to theslant range, the elevation, and the azimuth of the missile as measuredat the location of the missile tracking device. As shown in FIG. 1,however, certain advantageous modifications have been made in themissile tracking device to improve the performance thereof. Basicallythese modifications involve recognition of the fact that theantiaircraft missile presents an extremely difficult target for atracking radar. Accordingly it has been found desirable to employ aso-called radar beacon system rather than a conventional radar. For thispurpose pulses are radiated from a transmitter 24 at a radio frequency fThe missile as will be explained in greater detail hereinafter carries aresponder which is responsive to pulses of frequencies f and radiatedpulses of frequencies f These pulses are picked up by antenna 25 anddirected to a receiver 26 responsive to that radio frequency. Thereceiver 26 operates in a manner identical to that of target trackingradar 10 to provide the required output quantities as to azimuth andelevation for transmission over a connector 28 by way of a transferswitch 112 operated by an actuator 114 and parallax correction unit 31to a second predictor 30. The output derived from the range unit ofreceiver 26 and representative of the range to the missile is applieddirectly to parallax correction unit 31 by way of connector 29.

Normally and during all phases of an engagement other than the end game(the final increments of the engagement), transfer switch 112 remains inthe position shown in FIG. 1 whereby the output quantities from missiletracking radar 12 are applied after correction for parallax to predictor30 of computer 14.

Radar beacon systems of the type contemplated for use in the missiletracking system are well known in the art and are discussed in detail inRadar Beacons" by Roberts, Vol. 3 of The Radiation Laboratory Series,McGraw-I-Iill, 1947. Modification of the SCR-S 84 radar referred toabove as illustrative of the ground based missile tracking device, forthis type of performance may easily be accomplished merely by tuning ofthe receiver to frequency f rather than f,. The missile borne equipmentwill be considered hereinafter.

As assumed above the quantities applied to predictor 22 indicate thepresent position of target 18 in spherical coordinates with respect tothe location of the target tracking radar while those applied topredictor 30 represent similar information as to the position of themissile with reference to the location of the missile tracking radar.For ease in computation it is considered desirable to convert thisinformation into rectangular coordinates with the origin at the locationof the target tracking radar. Such coordinate conversion is well knownin the art and may be accomplished as described, for example, in U.S.Pat. No. 2,408,081 to Lovell et al. which issued, Sept. 24, 1946.Conveniently each predictor 22 and 30 includes a coordinate converteracting to convert input quantities to rectangular coordinates (X, Y andH where X and Y are orthogonal axes in the horizontal ground plane and His the vertical distance from the XY plane) with origins at thelocations of the respective tracking radars. The necessary offset orparallax corrections referred to above and required to convert the dataas to missile position to the coordinate system having its origin at thelocation of the target tracking radar may be set in manually in unit 31associated with predictor 30 as potentials of suitable polarity alongthe three rectangular coordinates. These corrections are constants andonce determined with adequate precision at the time at which the twotracking radars are emplaced, need not be changed unless the emplacementof the guidance equipment is changed. As pointed out above, however, thepresent invention permits substantial relaxation of the degree ofprecision required in these corrections. The errors in performanceresulting from parallax are not of great importance in the early phasesof an engagement. In the final phase control is switched according tothe invention to eliminate the parallax correction problem in itsentirety.

Computer 14 which includes predictors 22 and 30 acts on the basis of anassumed time of flight for the missile to reach a point of interceptionwith the target. Using this time of flight the future positions of thetarget and missile at the predicted time of interception are computedindependently. These positions are compared coordinate by coordinate tofind the predicted position error at the predicted time of interception.

The corrections in either or both the time of flight as- Y sume'd at theoutset and the course of the missile required to insure interception maybe determined from this information. Depending upon the nature of themissile either or both of these quantities may be altered to eliminatethe position error by the predicted time of interception. This generalphilosophy of computation is similar in certain respects to thatemployed in antiaircraft gun direction systems of the type disclosed inthe patent referred to above. In the computer herein disclosed, however,many of the operations are duplicated to provide information not only asto the target but also as to the missile since in the present case somemeasure of control remains after the missile is fired;

Assuming a time of flight. predictors 22 and 30 determine the futurepositions of both missile and target at the predicted time of intercept(T=0 where T is the time until interception). In each predictor thepresent position data is differentiated to obtain velocity and thisvelocity is multiplied by the time of flight to obtain future position.These operations are carried out in each instance in terms of componentsalong the orthogonal X, Y and H coordinates which are ground coordinateswith the origin at the location of the target tracking radar asindicated in FIG. 3 of the drawing.

As in the reference patent each predictor 22 and 30 provides threecoordinate output data representing the predicted position of the targetor the missile as the case may be when the predicted time of interceptoccurs. These quantities are identified in FIG. 1 as X Y and H for thetarget and X Y and H for the mis- I sile.

These quantities are applied to a comparator 32 in which are subtractedcoordinate by coordinate to obtain predicted position error quantities.The error outputs shown at the output of comparator 32 are E showing theposition error in the H direction, and E and E) showing the positionerror in the X and Y direc- H referred to above and the target aposition b similarly specified by the quantities X Y and H The totalposition error at T=0 for these assumptions is measured by the vector E.the components of which E E and E in the X, Y and H directionsrespectively are as shown in FIG. 3. It is apparent that thesequantities are measured with reference to a set of coordinates fixedwith respect to the target tracking radar and do not indicate directlywhat maneuvers must be performed by the missile to assure interception.It is necessary, therefore. to convert these error quantities intoquantities measured with respect to a frame of reference traveling withthe missile so that appropriate commands for missile guidance may beproduced.

The method of providing the required frame of reference traveling withthe missile and still identifiable at the location of the guidanceequipment will become apparent with reference to the diagram of FlG. 2.Here the missile 16 is shown as carrying a so-called free-free gyroscope34, the rotor of which is suspended in a conventional gimbal system. Theouter gimbal 36 is journalled for rotation about the longitudinal axisof the missile while the inner gimbal 38 is journalled in the outergimbal for rotation about an axis normal to the longitudinal axis of themissile. The gyroscope rotor is in turn journalled in the inner gimbaland spins about an axis normal to the inner gimbal axis. This third axisis hereinafter referred to as the gyro spin axis. In accordance withwell understood principles the gyroscope acts to maintain the gyro spinaxis G at a fixed orientation in space regardless of the maneuvers ofthe missile in which the gyroscope is mounted.

Conveniently, the gyro spin axis is initially oriented in the XY(horizontal) plane and, if possible, normal to the plane of the initialtrajectory to the target. This axis and the longitudinal axis of themissile define a reference plane attached to the missile andidentifiable at all times at the location of the guidance equipment aswill be pointed out below.

The initial orientation of the spin axis may be determined before themissile is launched and the heading of the missile at any time isdetermined as one of the necessary functions of predictor 30. It will berecalled that the rates of change of each of the X, Y and H quantitiesrepresenting the present position of the missile are determined in theprediction process. These components of the missile velocity may becombined to give a quantity proportional to the missile velocity in thedirection of the missile flight path. This combination of velocitycomponents is performed by missile heading resolver 40 which is aconventional coordinate resolution device of the type employed generallyin fire control computers. It is noted that the frame of referencetraveling with the missile is based upon the gyroscope spin axis and thelongitudinal axis of the missile while the information available to thecomputer includes the orientation of the gyro spin axis and the missilevelocity. It has been found that, because of the continuous control ofthe missile afforded by the command system of the invention, the missilevelocity vector may be considered to have the same orientation as thelongitudinal axis of the missile. Any error so introduced is within therange of correction of the command system. It will be understood,however, that if sufficient information is available as to theaerodynamic performance additional computation equipment may be providedto determine the orientation of the longitudinal axis of the missilefrom the available missile velocity information.

It now becomes necessary to steer the missile with respect to the frameof reference just considered. As shown in FIG. 2 the missile is providedwith two sets of paired steering fins 40 and 42 respectively. Fins 40are mounted on a shaft 44 normal to the longitudinal axis of the missileand fins 42 are mounted on a shaft 46 which is normal to both thelongitudinal axis of the missile and the shaft 44. Shafts 44 and 46 thusconstitute a pair of steering axes which will be referred to as the yawand pitch axes respectively and which control the orientation of themissile in two orthogonal planes. As a matter of convenience the planenormal to shaft 44 will be referred to as the yaw plane (which is thesame as the reference plane considered above) and that normal to shaft46 as the pitch plane (which includes the longitudinal axis of themissile and is normal to the reference plane). Fins 40 then constitutethe yaw steering fins and fins 42 the pitch steering fins. These finsare positioned in response to steering orders transmitted from computer14 to control the course of the missile after launching.

It is apparent that the course of the missile can be properly controlledonly if the steering axes represented by shafts 44 and 46 remainrespectively normal to and in the plane of reference considered aboveand defined by the gyro spin axis and the longitudinal axis of themissile. This condition requires roll stabilization of the missile. Forthis purpose the missile is equipped with paired cruciform tail fins 48and 50 and adjustable ailerons 52 are provided upon the trailing edgesof at least one pair of tail fins (48 in FIG. 2). Conveniently gyroscope34 is employed to control ailerons 52in such a way as to roll stabilizethe missilewith respect to the reference plane.

The outer gimbal 36 of the gyroscope 34 is coupled to the movable arm ofa potentiometer 54 through a shaft 56 which forms an extension of theouter gimbal axis. Potentiometer 54 comprises a source of error signalfor a servo system including an amplifier 56, a motor 58 and suitablelinkage 60, 62 whereby opposite deflections of upper and lower ailerons52 may be produced by motor 58. Although these elements, together withthe gyroscope 34 which constitutes the error detecting element, may beconnected to form any of a large number of known kinds of servo system,it may be assumed for the purposes of the present description that asimple direct current servo system is employed.

Potentiometer 54 is provided with a center-tapped winding to whichdirect current potentials are applied from a source such as a battery(not shown to avoid undue complexity in the drawing). The circuit is soarranged that no output is produced when shaft 44 is normal to thereference plane. Whenever the potentiometer arm is moved from thisnormal (null) position an output is developed, the amplitude andpolarity of which are indicative of the amount and direction of the rollof the missile from the desired position. After amplification thisoutput may control motor 58 as in the usual direct-current servo system,causing deflection of the ailerons in the proper direction to return themissile to the desired orientation as indicated by the null output frompotentiometer 54.

It will be understood that as a result of such roll stabilization fins40 are effective to produce steering forces in the yaw plane and theother pair of steering fins 42 act to produce steering forces in thepitch plane. Thus there is provided a set of orthogonal referencecoordinates traveling with the missile and comprising the yaw axis y,the pitch axis p and the missile heading h Further the orientation ofthis reference system in space is continuously determinable at thelocation of the ground guidance equipment.

It will be recognized that by the usual process of coordinate conversionthe total position error E shown in FIG. 3 may be expressed in terms ofcomponents along the reference axes traveling with the missile. Suchconversion may be accomplished in coordinate resolver 64, FIG. 1 asoutlined beginning page 279 of Electronic Analog Computers by Korn andKorn or in U.S. Pat. No. 2,658,674 to Darlington et al., Nov. 10, 1953which resulted from an application filed Feb. 13, 1945 particularly inFIG. 38 and the specification beginning at column 92 thereof. Thiscoordinate resolver accepts the three position error components E E andB shown in FIG. 3 and in addition accepts quantities from the output ofmissile heading resolver 40 indicating the orientation of missileheading axis in space and quantities representing the position of thegyro spin axis G which may be set into the coordinate resolver asconstants. The outputs of coordinate resolver 64 are E and E,,representing position errors in the yaw and pitch planes and measuredalong the pitch and yaw axes respectively and E, representing a positionerror along the missile path. These components are shown in FIG. 3 withrespect to the missile axes p, y and h,,,, (drawn with an origin atpoint c, the present position of the missile.)

E,, the predicted position error along the missile path is a measure ofthe adjustment which must be made in the time of flight (or thevelocity) of the missile to cause interception of the target. If it beassumed that the velocity of the missile is not subject to externalcontrol once the missile is launched, this correction must be made byvarying the time of flight originally assumed at the outset of thecomputation process and employed as one input to each of predictors 22and 30. This quantity is, therefore, applied to a time of flight servomechanism 60 and controls the setting of input quantities to the twopredictors possibly as a shaft rotation as in the predictors shown inU.S. Pat. No. 2,408,081, referred to above. The quantity controllingpredictor 22 is applied directly thereto. However, the correspondingquantity for predictor 30 which is associated with the missile sectionof the computer is modified in accordance with the ballisticcharacteristics of the missile before application to predictor 30. Suchmodifications are accomplished by apparatus 62 wherein appropriatechanges are made in the value of time of flight. These changes areordinarily accomplished by adding both fixed and variable components tothat corresponding to the time of flight as shown for example in FIG. 8Aof U.S. Pat. No. 2,408,081 to which reference has been made above.

It will be understood from the above that the necessary functions forthe continuous prediction process performed by the computer are providedby way of the time of flight servo mechanism and the target and missiletracking radars. The pitch and yaw error outputs E and E from coordinateresolver 58 are employed for the generation of steering orders for themissile. These quantities depend of course upon the time of flight fedback to predictors 22 and 30 as discussed above. As a matter ofconvenience in control of the missile it has been found desirable toconvert these position orders into acceleration orders, i.e., thequantities representing the position errors measured along the yaw andpitch axes are converted into lateral accelerations in the pitch and yawplanes respectively such that the missile will be at the point ofpredicted interception at the predicted time of interception. For thegeneration of such orders the quantities E and E are appliedrespectively to dividers and 72 in which each is divided twice by thetime of flight produced as the output of unit 66. The yaw and pitchorders appearing at the outputs of dividers 70 and 72 respectively arethus proportional to the accelerations required to cause the missile toreach the predicted point of interception at the predicted time ofinterception. These orders may be transmitted to the missile by anyconvenient means, for example, by a high frequency radio communicationchannel.

Alternatively and as shown in FIG. 1 the acceleration orders for themissile are transmitted by modulation of the repetition rate of themissile tracking pulse transmitter. The necessary control quantities maybe trans mitted on a time division basis or any other convenient basisby the action of a modulator 74 associated with transmitter 24. Varioussystems of signaling over the radar beam are described in Section ll.2of Radar Beacons," Vol. 3 of the Radiation Laboratories Series.According to one such system the two control signals are transmitted asaudio frequency signals frequency modulated upon the pulses from thetrack transmitter. Either one or both of the frequencies can thus betransmitted depending upon the steering orders required at a particulartime, the modulating wave comprising either one or the sum of the audiofrequencies.

Also transmitted to the missile and conveniently by interruption of allmodulation upon the radar beam is the so-called burst order which at atime related to the predicted time of intercept causes the warhead ofthe missile to explode. Ordinarily this order is transmitted a fewmicroseconds prior to the time when T=0.

The remaining equipment carried aboard the missile may now beconsidered. As has been stated above the missile carries a transponderresponsive to pulser from the ground base missile tracking equipment.This transponder includes a receiver 68 tuned to frequency f associatedwith antenna 70, and a microwave pulse transmitter 72 which is triggeredby the output of receiver 68 and which radiates pulses of frequency ffrom antenna 74. These pulses when received at the location of theguidance equipment permit tracking of the missile.

Radio receiver 68 also serves to receive the various orders transmittedfrom the guidance equipment and intended to control the steering fins ofthe missile and the bursting of the warhead at appropriate times.Depending upon the nature of the modulation employed to transmit theseorders this portion of receiver 68 may include a frequency modulationdemodulator, pulse position demodulators or decoders or a frequencydivision multiplex receiver wherein the various order channels aredistinguished upon a frequency basis. In any event the receiver isdesigned with reference to the particular transmitter 24, FIG. 1,employed in the guidance equipment and produces three output signalscorresponding respectively to the pitch and yaw acceleration orders andthe warhead burst order.

As shown in FIG. 2 each pair of missile steering fins is driven by anelectric motor, in response to the appropriate orders occurring at theoutput of receiver 68. Thus a motor 76 is geared to the shaft 44 uponwhich steering fins 40 are mounted and a motor 80 is geared to thecorresponding shaft 46 upon which steering fins 42 are mounted. It willbe recalled that the missile is to be made responsive to accelerationorders. Accordingly for each set of steering fins the appropriate orderappearing at the output of radio receiver 68 is applied to an amplifierto which is also applied an output of an accelerometer. These twoquantities are applied in opposition and the motor is driven from theoutput of the amplifier until the accelerometer indicates that thedesired angular acceleration has been introduced. When such a conditionoccurs the output of the amplifier is reduced to zero and the motorstops. If the angular acceleration increases. the output of theaccelerometer exceeds the order output of the receiver and the motor isdriven in the appropriate direction to return the acceleration to therequired value. The yaw steering fins 40 are thus controlled by theoutput of the comparison amplifier 84 to which is applied the output ofan accelerometer 86 oriented in the missile to indicate accelerationsabout the yaw axis. Similarly the pitch steering fins 42 are controlledby a comparison amplifier 88 to which is applied the pitch order outputof the receiver and the output of an accelerometer 90 oriented in themissile to detect accelerations about the pitch axis.

The remaining equipment in the missile includes a warhead 92 furnishedwith an appropriate detonator which may be actuated by an electricimpulse known as the burst order received from the third output ofreceiver 68 and applied to the warhead over lead 94.

In the operation of the missile system as thus far described the missileis launched and guided toward interception with the target in the samemanner as that disclosed in the copending application, Ser. No. 449,396referred to above. However, and in accordance with the present inventionthe missile tracking function is transferred from the missile trackingradar 12 to the target tracking radar which is modified as describedabove so that during the final phases of an engagement,

referred to above in some instances as the end game,

both target and missile are tracked in elevation and azimuth by the sameradar, in this instance target tracking radar 10. Under thesecircumstances those data as to position of both target and missile whichare most sensitive to error are referred to the same reference systemand such matters as parallax between the two tracking radars and theaccurate bore-sighting of both as required to insure appropriatecorrections are substantially eliminated. Conveniently the time in whichtransfer of control from the missile tracking radar to the targettracking radar is effected is determined by the reduction of the angularseparation of the tracking beams of target tracking radar l0 and missiletracking radar 12 to a predetermined small amount. Although thiscriteria for switching is considered the most desirable because of thefact that in the normal engagement the v missile and target tend to bein approximate alignment in the final phases of the engagement, othercriteria for switching may well be employed. For example, switching maybe effected when the time to go to predicted interception has beenreduced to a predetermined value. In any event, switching should notoccur until the missile has approached sufficiently near to the targetto be illuminated by the beam of the target tracking radar.

In any event information as to the reduction of the switching criteriato a predetermined value is transmitted over leads 116 and 118 fromtarget tracking radar l0 and missile tracking radar 12, respectively, toa control circuit 120. When the predetermined criteria has been realizedcontrol circuit 120 becomes effective to operate actuator 114 switchingthe input of predictor 30 from the output of parallax unit 31 associatedwith lead 38 from missile tracking radar 10 to lead 108 from receiver104 of the target tracking radar. It will be noted that this transfereliminates the parallax correction unit 31 from the elevation andaximuth inputs to predictor 30 since such correction is no longerrequired. From this time onwards during the engagement both target andmissile are illuminated by the beam from target tracking radar l0 andthe automatic tracking equipment of this radar maintains it in suchorientation as to keep both target and missile within the beam.

The beacon signals from missile 16 may thus be received by receiver 104associated with the target tracking radar and from these receivedsignals may be derived error quantities representative of the failure ofthe target tracking antenna to track the missile exactly in elevationand azimuth. In the conventional tracking radar, these error signals areemployed as inputs to servo devices arranged to orient the antenna andadjust the range unit in such a way as to reduce the errors in trackingeffectively to zero. Here, however, antenna orientation and range unitadjustment are and remain under control of the target tracking radar andare effected in response to the error signals from receiver 104. Theerror signals AA,,,, and A5, representing the azimuth and elevationerrors in the tracking of the missile by the target tracking radar aretherefore applied to an adder 110. This device is of conventional typeand includes three channels in which AA, and AE are algebraically addedto A and E respectively as derived from servo unit 106. The output ofadder 110 as applied to transmitting channel 108 includes separatequantities representing A and E of the missile position data as requiredby predictor of the computer. Missile range data for application topredictor 30 continues to be derived from the range unit of missiletracking radar 12 and is applied by way of parallax correction unit 31as noted previously. These quantities representing angular position areapplied to predictor 30 in preference to those normally supplied bymissile tracking radar l2 and the computer treats them in the samemanner as though they were derived from missile tracking radar 12 andthe remainder of the system operates exactly as previously described.The single reference for both missile and target position data is thusachieved for the azimuth and elevation quantities and the need foraccurate parallax and bore-sight adjustments is eased during that phaseof the engagement for which these quantities become of the greatestimportance.

The control system by means of which the approach of angular separationof the two tracking beams to within the predetermined small quantity,for example 1 in both elevation and azimuth, may be detected will now bedescribed with reference to FIG. 4 of the drawings. Here the antennaplatforms 122 and 124 of target and missile tracking radars respectivelyare shown schematically. These units may include any combination ofsynchros, converters, motors, etc., as required according to well knownprinciples, to provide an output which is the differential between twoinputs. The exact nature of the units will be dependent upon the natureof the inputs and outputs which may be electrical quantities ormechanical shaft rotations as indicated in FIG. 4. Each of theseplatforms will be understood to provide for rotation of the antennaproper in elevation and azimuth. In the control circuitry now to beconsidered the azimuth shaft of missile tracking radar plat form 124 isconnected to a synchro transmitter 126 while the azimuth shaft isconnected to a similar transmitter 128. In a similar manner theelevation shaft of target tracking radar platform 122 is connected asone input to a differential synchro unit 130 and the azimuth shaft isconnected as one input to a differential synchro unit 132. For ease ofillustration, the various components such as synchros are shown adjacentto the antenna platforms to which they are linked. It will be understoodthat with the exception of antenna platforms 122 and 124, all of theelements shown in FIG. 4 are a part of the control circuit shownschematically in FIG. 1. The electrical output of elevation transmitter126 is applied as the second input as the differential synchro and theelectrical output of azimuth 128 is applied as the second input todifferential synchro 132. The output shaft of differential synchro 130is mechanically connected to a switch 134 and the corresponding shaft ofdifferential synchro 132 is connected to a switch 136. Switches 132 and136 are simple single pole switches through which may be completed aseries circuit including a source of potential such as a battery 138 andactuator 114 of FIG. 1.

The mechanical output of each of differential synchros 130 and 132constitutes a measure of the angular separation about the associatedaxes of radar platforms 122 and 124 thus switch 134 is closed wheneverthe two elevation shafts approach within 1 of angular separation andswitch 136 is closed whenever the azimuth shafts approach within thesame limit. When both switches are closed the requirement for switchingof the control from the missile tracking radar to the target trackingradar is made and actuator 114 closes transfer switch 112 and therequired transfer of control is effected.

What is claimed is:

1. In an antiaircraft system, target tracking means for continuouslyestablishing the position of a target aircraft in space, missiletracking means for similarly establishing the position of a missile,computing means responsive to position information from both said targetand missile tracking means to produce control signals for said missileto insure interception of the target thereby and means for substitutingthe target tracking means for said missile tracking means as a source ofmissile position information for said computing means during the finalphases of an engagementv 2. In an antiaircraft system a tracking radarfor establishing the position of a target aircraft in space, a secondtracking radar for similarly establishing the position of a missilelaunched against said target aircraft, computing means responsive toposition information from both of said radars to produce control signalsfor said missile to insure interception of the target aircraft thereby,means for detecting the approach of the missile to within apredetermined separation from said target aircraft and means responsiveto said detecting means for substituting the target tracking radar forsaid missile tracking radar as a source of missile position information.

3. In an antiaircraft system target tracking means for continuouslyestablishing the azimuth, elevation and range to a target aircraft,missile tracking means for similarly establishing the position of amissile, computing means responsive to the position information fromboth target and missile tracking means to control the path of saidmissile to insure interception of the target thereby and means forsubstituting the target tracking means for said missile tracking meansas the source of missile azimuth and elevation information during thefinal phases of an engagement.

4. In an antiaircraft system a missile tracking radar operating at afirst frequency, a target tracking radar comprising a transmitteroperated at a second frequency and receiving means capable of separateresponse to echo signals of both said first and second fre quencies,computing means normally responsive to tar- 13 get and missile positioninformation derived from said first and second radars respectively tocontrol the path of said missile to insure interception of the targetthereby and means operative during the final phase of an engagement tosubstitute as a source of missile position information for said computerthe output of the i for determining the angular separation in bothelevation and azimuth of the tracking beams of said radars and meansoperative when both of said angular separations are reduced belowpredetermined values for substituting the target tracking radar for saidmissile tracking radar as a source of missile position information forapplication to said computing means.

6. Apparatus for controlling a guided missile relative to a targetcomprising missile tracking radar apparatus,

' target tracking radar apparatus and computer apparatus connected withsaid respective missile tracking and target tracking radar apparatus,said target tracking radar apparatus comprising means independentlyreceiving radio signals from said missile and said target forsimultaneously supplying said computer apparatus with informationsignals in accordance with the angular position of said missile and therange and angular position of said target.

7. In an antiaircraft system target tracking means for continuouslyestablishing the position of a target aircraft in space, missiletracking means independent thereof for similarly establishing theposition of a missile launched against said target aircraft, a computer.means for applying the output of said target tracking means to saidcomputer, means for applying the output of said missile tracking meansto said computer after correction for parallax between said missile andtarget tracking means, said computer being arranged to produce ordersfor said missile to insure interception of the target aircraft by saidmissile and means operative when the target aircraft and missileapproach one another to within a predetermined separation forsubstituting the output of said target tracking means as a source ofmissile position information in place of the missile tracking means andthe means for correcting for parallax to the information for which suchsubstitution is made

1. In an antiaircraft system, target tracking means for continuously establishing the position of a target aircraft in space, missile tracking means for similarly establishing the position of a missile, computing means responsive to position information from both said target and missile tracking means to produce control signals for said missile to insure interception of the target thereby and means for substituting the target tracking means for said missile tracking means as a source of missile position information for said computing means during the final phases of an engagement.
 2. In an antiaircraft system a tracking radar for establishing the position of a target aircraft in space, a second tracking radar for similarly establishing the position of a missile launched against said target aircraft, computing means responsive to position information from both of said radars to produce control signals for said missile to insure interception of the target aircraft thereby, means for detecting the approach of the missile to within a predetermined separation from said target aircraft and means responsive to said detecting means for substituting the target tracking radar for said missile tracking radar as a source of missile position information.
 3. In an antiaircraft system target tracking means for continuously establishing the azimuth, elevation and range to a target aircraft, missile tracking means for similarly establishing the position of a missile, computing means responsive to the position information from both target and missile tracking means to control the path of said missile to insure interception of the target thereby and means for substituting the target tracking means for said missile tracking means as the source of missile azimuth and elevation information during the final phases of an engagement.
 4. In an antiaircraft system a missile tracking radar opErating at a first frequency, a target tracking radar comprising a transmitter operated at a second frequency and receiving means capable of separate response to echo signals of both said first and second frequencies, computing means normally responsive to target and missile position information derived from said first and second radars respectively to control the path of said missile to insure interception of the target thereby and means operative during the final phase of an engagement to substitute as a source of missile position information for said computer the output of the target tracking receiving means produced in response to echoes of said second frequency in place of the output of the missile tracking radar.
 5. In an antiaircraft system a target tracking radar for continuously establishing the position of a target aircraft in space, a missile tracking radar for similarly establishing the position of a missile launched against said target aircraft, computing means responsive to position information from both said target and missile tracking radars to produce control signals for said missile to insure interception of the target aircraft thereby, means for determining the angular separation in both elevation and azimuth of the tracking beams of said radars and means operative when both of said angular separations are reduced below predetermined values for substituting the target tracking radar for said missile tracking radar as a source of missile position information for application to said computing means.
 6. Apparatus for controlling a guided missile relative to a target comprising missile tracking radar apparatus, target tracking radar apparatus and computer apparatus connected with said respective missile tracking and target tracking radar apparatus, said target tracking radar apparatus comprising means independently receiving radio signals from said missile and said target for simultaneously supplying said computer apparatus with information signals in accordance with the angular position of said missile and the range and angular position of said target.
 7. In an antiaircraft system target tracking means for continuously establishing the position of a target aircraft in space, missile tracking means independent thereof for similarly establishing the position of a missile launched against said target aircraft, a computer, means for applying the output of said target tracking means to said computer, means for applying the output of said missile tracking means to said computer after correction for parallax between said missile and target tracking means, said computer being arranged to produce orders for said missile to insure interception of the target aircraft by said missile and means operative when the target aircraft and missile approach one another to within a predetermined separation for substituting the output of said target tracking means as a source of missile position information in place of the missile tracking means and the means for correcting for parallax as to the information for which such substitution is made. 